Rotorcraft autopilot system, components and methods

ABSTRACT

An autopilot actuator includes first and second motors each including a rotatable motor output shaft such that either one or both of the motors can drive an actuator output shaft. An autopilot main unit enclosure is removably mounted to the helicopter proximate to a cyclic control and commonly houses autopilot actuators as well as main autopilot electronics. A cyclic vibration isolator is removably supported by an actuator shaft for co-rotation and coupled to the cyclic control to attenuate a cyclic vibration frequency at the actuator shaft while output rotations of the actuator shaft below a resonant frequency are coupled to the cyclic control. A force limited link includes first and second ends and a variable length between. The force limited link having a relaxed length when less than an unseating force is applied and the variable length changes when an applied force exceeds the unseating force to permit pilot override.

RELATED APPLICATIONS

This application is a divisional application of copending U.S. patentapplication Ser. No. 13/763,590 filed on Feb. 8, 2013, the disclosure ofwhich is incorporated herein by reference and which claims priority fromU.S. Provisional Patent Application Ser. No. 61/597,555; U.S.Provisional Patent Application Ser. No. 61/597,570; and U.S. ProvisionalPatent Application Ser. No. 61/597,581, each of which was filed on Feb.10, 2012 and all of which are hereby incorporated by reference in theirentirety.

BACKGROUND

The present application is generally related to flight control systemsand, more particularly, to a rotorcraft autopilot and associatedmethods.

A helicopter is inherently unstable, generally requiring that the pilotmaintain a constant interaction with the cyclic control using one hand.Even a momentary release of the cyclic can result in the cyclic orcontrol stick “flopping over”, accompanied by a loss of control of thehelicopter. This is particularly inconvenient when the pilot has a needto engage in hands-free activities such as, for example, adjusting aheadset or referring to a hardcopy of a map. Further, the need toconstantly control the cyclic can result in pilot fatigue.

Traditional autopilots can provide benefits which include allowing thepilot release the cyclic to engage in hands-free tasks, as well asreducing pilot fatigue. Applicants recognize, however, that the cost ofa traditional helicopter autopilot can be prohibitive. For example, thecost can be so significant in comparison to the cost of the helicopteritself that autopilots are uncommon in light helicopters.

The foregoing examples of the related art and limitations relatedtherewith are intended to be illustrative and not exclusive. Otherlimitations of the related art will become apparent to those of skill inthe art upon a reading of the specification and a study of the drawings.

SUMMARY

The following embodiments and aspects thereof are described andillustrated in conjunction with systems, tools and methods which aremeant to be exemplary and illustrative, not limiting in scope. Invarious embodiments, one or more of the above-described problems havebeen reduced or eliminated, while other embodiments are directed toother improvements.

Generally, an autopilot system for a helicopter associated componentsand methods are described. In one aspect of the disclosure and as partof an autopilot for providing automatic control of a helicopter byactuating one or more flight controls, an actuator is configured toinclude a redundant pair of first and second motors each including arotatable motor output shaft. A gear arrangement includes an actuatoroutput shaft for operative coupling to the flight controls and isconfigured to engage the output shaft of each one of the first andsecond motors for operation at least in (i) a first mode in which boththe first and second motors contribute to rotation of the actuatoroutput shaft, (ii) a second mode in which first motor rotates theactuator output shaft due to a failure of the second motor and (iii) athird mode in which the second motor rotates the output shaft due to afailure of the first motor.

In another aspect of the disclosure, an autopilot display unit ismounted in an instrument console of the helicopter at least fordisplaying autopilot flight data to a pilot of the helicopter. Anautopilot main unit enclosure is removably mounted to the helicopterproximate to a cyclic control of the helicopter and which defines a mainunit interior. A set of actuators is supported within the main unitinterior including a pitch actuator having a pitch actuator output shaftand a roll actuator having a roll actuator output shaft such that nomore than the pitch actuator shaft and the roll actuator shaft of theset of actuators extend at least partially outward from the autopilotmain unit enclosure for providing mechanical control forces to thecyclic control of the helicopter. A main unit electronics section issupported within the main unit interior and is in electricalcommunication with the autopilot display unit and with the set ofactuators for providing electrical control signals to the actuators suchthat the main unit electronics section and the set of actuators arecommonly housed within the main unit interior.

In still another aspect of the disclosure, an autopilot system isconfigured for automated control of a helicopter by actuating a cycliccontrol of the helicopter which cyclic control is subject to a cyclicvibration frequency responsive to rotation of a rotor of the helicopter.A cyclic vibration isolator is removably supported by the actuator shaftfor co-rotation therewith and coupled to the cyclic control of thehelicopter to exhibit a resonant frequency that is at leastapproximately matched to the cyclic vibration frequency for movement ofthe cyclic control relative to the actuator shaft such that the cyclicvibration frequency is attenuated at the actuator shaft and outputrotations of the actuator shaft below the resonant frequency are coupledto the cyclic control.

In yet another aspect of the disclosure, an autopilot system isconfigured for automated control of a helicopter by actuating a cycliccontrol of the helicopter which cyclic control is subject to a cyclicvibration responsive to rotation of a rotor of the helicopter. A cyclicvibration isolator includes a control arm that is removably attached tothe actuator shaft to support the cyclic vibration isolator such thatthe control arm co-rotates with the actuator shaft. An output arm iscoupled to the cyclic control of the helicopter and therefore subject tothe cyclic vibration. A resilient arrangement is captured between thecontrol arm and the output arm such that the output arm oscillatesresponsive to the cyclic vibration and relative to the control arm tomechanically isolate the actuator shaft from the cyclic vibration whiletransferring rotational actuation motions of the actuator shaft to theoutput arm to thereby transfer the rotational actuation motions to thecyclic control for autopilot actuation of the cyclic control.

In a continuing aspect of the disclosure, an autopilot system isconfigured for automated control of a helicopter by driving an actuatorhaving an actuator shaft to actuate a cyclic control of the helicopter.As part of the autopilot, a force limited link includes a first end, asecond end and a variable length therebetween oriented along anelongation axis with the first end coupled to the actuator shaft and thesecond end coupled to the cyclic control. The force limited link havinga relaxed length as the variable length between the first and secondends when less than an unseating force is applied to the first andsecond ends along the elongation axis to provide compliant movement ofthe cyclic control responsive to the actuator and configured such thatthe variable length changes from the relaxed length responsive to anexternal force that is applied to the first and second ends along theelongation axis which is equal to or greater than the unseating force toprovide for a pilot override of actuator cyclic control.

In a further aspect of the disclosure, an autopilot system is configuredfor automated control of a helicopter by providing electrical commandsignals to an actuator to actuate a cyclic control of the helicopterwhich cyclic control is subject to a cyclic vibration responsive torotation of a rotor of the helicopter. An autopilot linkage includes aforce limited link to provide a pilot override of cyclic control by theactuator. The force limited link having a first end, a second end and avariable length therebetween oriented along an elongation axis with thefirst end coupled to the cyclic control. The force limited link having arelaxed length as the variable length between the first and second endswhen less than an unseating force is applied to the first and secondends along the elongation axis to provide compliant movement of thesecond end with the first end and configured such that the variablelength changes from the relaxed length responsive to an external forcethat is applied to the first and second ends along the elongation axiswhich is equal to or greater than the unseating force for pilotoverride. A cyclic vibration isolator is removably supported by theactuator shaft for co-rotation therewith and connected to the second endof the force limited link such that the cyclic vibration is coupled tothe cyclic vibration isolator by the force limited link and the cyclicvibration isolator is configured to exhibit a resonant frequency that isat least approximately matched to a cyclic vibration frequency of thecyclic vibration to allow movement of the second end of the forcelimited link relative to the actuator shaft at the cyclic vibrationfrequency while isolating the actuator shaft from the cyclic vibrationfrequency and for transferring output rotations of the actuator shaftbelow the resonant frequency to the force limited link for transfer tothe cyclic control.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments are illustrated in referenced figures of the drawings. It isintended that the embodiments and figures disclosed herein are to beillustrative rather than limiting.

FIG. 1 is a diagrammatic perspective, partial view of a helicopterincluding components of an autopilot according to the presentdisclosure.

FIG. 2 is an overhead diagrammatic perspective, partial view of thehelicopter of FIG. 1, shown here to illustrate further details withrespect to components of the autopilot system.

FIG. 3 is a diagrammatic, perspective partial view of an embodiment ofan actuator and an embodiment of a force limited link that can serve ascomponents of the autopilot of the present disclosure.

FIG. 4 is a diagrammatic, perspective view of an embodiment of a geardrive arrangement that can form part of the actuator of FIG. 3 alongwith a redundant pair of actuator drive motors.

FIG. 5 is a block diagram that illustrates an embodiment of theautopilot of the present disclosure.

FIGS. 6 and 7 are diagrammatic views, in perspective, of an embodimentof a main unit enclosure and associated components.

FIG. 8 is another diagrammatic view, in perspective, of the main unitenclosure of FIGS. 7 and 8, shown here to illustrate further details ofits structure.

FIGS. 9 and 10 are diagrammatic views, taken from differingperspectives, showing the main unit enclosure and actuator linkages inan installed condition in relation to a cyclic control of thehelicopter.

FIG. 11 is an assembly view, in perspective, of an embodiment of acyclic vibration isolator according to the present disclosure.

FIG. 12 is a diagrammatic view, in perspective, of a control arm thatforms part of the vibration isolator of FIG. 11.

FIG. 13 is a diagrammatic, partial view in perspective of the vibrationisolator of FIG. 11, shown in a way that reveals further details of itsinternal structure.

FIG. 14 is a diagrammatic view, in perspective, of an embodiment of aforce limited link of the present disclosure.

FIG. 15 is a diagrammatic, exploded view in perspective of the forcelimited link of FIG. 14.

FIGS. 16-18 are diagrammatic, cutaway views, in perspective, of theforce limited link of FIG. 14 shown in respective ones of relaxed,compressed and extended operational states.

DETAILED DESCRIPTION

The following description is presented to enable one of ordinary skillin the art to make and use the invention and is provided in the contextof a patent application and its requirements. Various modifications tothe described embodiments will be readily apparent to those skilled inthe art and the generic principles taught herein may be applied to otherembodiments. Thus, the present invention is not intended to be limitedto the embodiments shown, but is to be accorded the widest scopeconsistent with the principles and features described herein includingmodifications and equivalents. It is noted that the drawings may not beto scale and may be diagrammatic in nature in a way that is thought tobest illustrate features of interest. Descriptive terminology may beadopted for purposes of enhancing the reader's understanding, withrespect to the various views provided in the figures, and is in no wayintended as being limiting.

FIG. 1 is a perspective, partial view of a helicopter 10, shown here forpurposes of illustrating various components of an embodiment of anautopilot system 12 in relation to the helicopter. It should beappreciated that much of the physical structure of the helicopter itselfhas been rendered as invisible in FIG. 1 for purposes of illustrativeclarity, however, it is understood that this structure is present. Theautopilot of the present disclosure is electromechanical and can provideflight control of a helicopter without requiring a hydraulic flightcontrol system. The helicopter can be, by way of non-limiting example, aRobinson R22 helicopter. The teachings that are brought to light herein,however, can readily be adapted for use with any suitable helicopter,either currently available or yet to be developed. For example, theautopilot of the present disclosure can be used with helicopters havinghydraulic cyclic assistance, with or without the hydraulics functioning.

Helicopter 10 includes a stick or cyclic 14 having a control handle orgrip 18 that is configured for engagement with the hand of a pilot. Aswill be appreciated by one of ordinary skill in the art, stick 14 can bemoved fore and aft (toward and away from an instrument console 20) tocontrol pitch of the helicopter and transversely for purposes ofcontrolling roll of the helicopter in a coordinated manner to producecontrolled flight. Additional control inputs are provided by the pilotvia a pair of pedals in order to control the yaw orientation of thehelicopter by changing the pitch of a tail rotor. It is noted that theseyaw orientation control components have not been shown for purposes ofillustrative clarity but are understood to be present. In an embodiment,the pilot also remains in control of the collective of the helicopter aswell as the throttle settings. The autopilot of the present disclosure,however, can exert full control authority over stick 14 by moving thestick in any direction to the limits of its travel under appropriatecircumstances. Stick 14 passes below a deck 24 of the helicopter andengages pitch and roll linkages of the helicopter in a manner that isfamiliar to one of ordinary skill in the art so as to control cyclicactuation of the main rotor of the helicopter. In particular, a torquetube 25 a transfers roll actuations while a control rod 25 b transferspitch actuations. The term “cyclic” refers to the variation in pitch ofthe rotor blades of the helicopter on a per revolution basis. In thisregard, cyclic control can refer to manipulation of the stick or thestick itself can be referred to as the cyclic. An autopilot displayprocessor unit (ADPU) 28 can be mounted in instrument console 20 toprovide indications to the pilot as well as to provide processingcapability and other capabilities. It is noted that the ADPU is alsoshown in a further enlarged, inset view.

The cyclic, in particular, handle 18 includes a Switch Module Assembly26 that can be mounted as shown. Details of handle 18 are shown in afurther enlarged inset view. The switch module can contain switchesincluding an engage/disengage switch 29 a and a trim/mode “top-hat”switch 29 b (4-way, in the present embodiment). The top-hat switchallows the pilot to trim the course, speed and altitude. There can be atime-out feature in the autopilot processor which prevents switch faultsor wiring faults from causing continuous trimming. The mode switch canselect and deselect altitude, speed, hover or position hold modes basedon current flight conditions.

Still referring to FIG. 1, autopilot 12 implements cyclic controlthrough a number of component assemblies that are appropriately locatedon the helicopter. A main autopilot unit 30 is located below the maindeck of the helicopter. In the present embodiment, main unit 30 includesan L-shaped enclosure 31 that supports electronics as well as a pitchcontrol linkage 32 a and a roll control linkage 32 b, which may bereferred to generally or collectively by the reference number 32. Eachof these linkages includes an actuator that is located within the mainunit enclosure, as will be further described. A distal end of each ofthe linkages engages the lowermost end of stick 14 to implement what isknown as a parallel control system. In this regard, it should beappreciated that the original cyclic control linkages of helicopter 10between stick 14 and the rotor remain intact. That is, inputs from thehelicopter pilot as well as the autopilot are input directly to thestick. Details with respect to the pitch and roll control linkagesprovide for a parallel control input arrangement. A series typeautopilot control system, in contrast, requires breaking the originalcyclic control linkages of the helicopter between the stick and rotorsuch that the autopilot actuators can be inserted into the break. Itshould be appreciated that the teachings herein can readily be adaptedto a series control input embodiment.

Turning to FIG. 2, components of the helicopter and autopilot are shownin an overhead perspective view. In this view, a pitch actuator 60 a anda roll actuator 60 b (which may be referred to generally or collectivelyby the reference number 60) can be seen within L-shaped enclosure 31with the lid of the enclosure rendered transparent. Main unitelectronics 66 are located within the enclosure and are suitablyelectrically interfaced (not shown) both externally and to theactuators. It is noted that additional details with respect to asuitable embodiment of main unit electronics 66 have been described incopending U.S. patent application Ser. No. 13/763,574 (Attorney DocketNo. HTK-2), which is incorporated herein by reference in its entirety.

Referring to FIG. 3, an embodiment of actuator 60 that can be used forthe pitch and roll actuators throughout this disclosure is seen in aperspective view installed within enclosure 31 and connected to acontrol linkage 32. Each actuator includes a housing 82 having a geararrangement, yet to be illustrated, within the housing, dual motorsMotor A and Motor B, and a clutch arrangement 84 for selectivelyengaging and disengaging the motors to rotate an output shaft which isnot visible on the opposite side of housing 82. The latter can beformed, for example, from stainless steel. As will be seen, the geararrangement allows motors A and B to simultaneously drive the outputshaft or either one of the motors to individually drive the outputshaft. In the present embodiment, motors A and B are brushless DC motorshaving a Y stator winding configuration which requires coordinatedinputs to drive the motor phases in a particular sequence that iswell-known. As such, the motors cannot runaway under their own power.The motors include Hall effect sensors that are used for purposes oftiming electrical drive pulses to the stator of the motor. Furtherdetails with respect to the motors and related drive considerations areprovided at one or more appropriate points hereinafter.

FIG. 4 illustrates an embodiment of a gear drive arrangement 100 thatcan be used in the actuator of FIG. 3. Initially, it is noted that thegear drive arrangement is a multi-stage reduction drive, for example, onthe order of about 1750:1. Also, teeth have not been illustrated on anumber of the gears to be described, but are understood to be present.Other embodiments may not require gears with teeth. Motors A and B haveoutput shafts 90 a and 90 b, respectively, supporting gears that engagea gear 102 on a first shaft 104. An opposing end of shaft 104 supports asmaller gear 106 that drives a gear 110 that is supported on a secondshaft 112 which also supports a smaller gear 114 (partially hidden inthe view of the figure). It is noted that shaft 112 can comprise aclutch shaft that can move laterally to selectively engage or disengagethe actuator motors from the remaining gears of the gear drive. Asuitable clutch arrangement is described, for example, in U.S. Pat. No.7,954,614 which is incorporated by reference. The clutch arrangementrelies upon movement of the clutch shaft along its elongation axis byusing a permanent magnet that is mounted on a distal end of the shaft.Clutch actuator 113 (FIG. 3) can selectively move (for example, rotate)another permanent magnet in relation to the clutch shaft mountedpermanent magnet such that the clutch shaft is magnetically biased tomove between an engaged position and a disengaged position. The clutchshaft remains in a current operational position despite a power failure.Gear 114, in turn, selectively drives a gear 120 that is supported on athird shaft 122. The latter also supports a smaller gear 124 that drivesa gear 130 that is supported on a forth shaft 132. The forth shaft, inturn, supports a smaller gear 134 which is arranged to rotate an outputgear 140 that is supported on an output shaft 142 of the actuator. Theoutput gear is configured to provide sufficient rotation to move stick14 through its full range of motion. In an embodiment, the actuators ofthe present disclosure are sufficiently robust, in terms of thegenerated level of actuation force, so as to be capable of controllingthe cyclic of a hydraulically equipped helicopter using a failedhydraulic system. For example, actuator 60 is capable of applying forcesof at least 100 pounds to the cyclic. While the present embodiment hasbeen designed to provide actuation forces at this level using anavailable output torque of up to 200 inch-pounds, it should beappreciated that in another embodiment, significantly higher force canbe provided, for example, by reducing the length of the actuator controlarm. As will be further described, the actuator forces are applied tothe bottom of the cyclic whereas pilot forces are applied to the top ofthe cyclic. Accordingly, the pilot is provided with a mechanicaladvantage due to the different lever-arm lengths. On the R22 helicopter,the mechanical advantage that the pilot has at the top of the stickcompared to the bottom of the stick where the actuators are attached isroughly 7:1. In such a case, an actuator applied force of 100 pounds isequivalent to about 14 pounds of pilot applied force. Similarly, whilethe actuator can generate very large forces, the force-limited link thatis described below generally will not be embodied to transmit forces ofsuch a magnitude through to the base of the cyclic, however, a muchstiffer force-limited link embodiment can be installed, if so desired.

In an embodiment, the actuator can be configured with a gear ratio ofapproximately 1720:1, however, a wide range of different gear ratios maybe found to be suitable. It should be appreciated that for a gear ratioof 1720:1, one revolution of the motor rotates the actuator output shaftby only about 0.2 degrees. In and by itself, this resolution can besufficient for monitoring the actuator output position. For example,rotation of the motor shaft can be detected using a magnet that ismounted on the shaft, as is familiar to one having ordinary skill in theart. In an embodiment, as described in the above incorporated U.S.patent application Ser. No. 13/763,574 (Attorney Docket No. HTK-2), Hallsensor data from the motors can be used to determine the incrementalposition of the actuator output shaft of each actuator. In this regard,each actuator motor includes 3 Hall sensors. The Hall sensor pulses canact like an incremental up/down counter. The position of the outputshaft relative to a reference location can be tracked constantly. Forexample, a zero reference position of the actuator output shaft can bedefined when the actuator is engaged via clutch 84. Such zero referenceposition tracking can be used for certain failures wherein the bestapproach resides in restoring the actuator arms/shafts to their averagedpositions prior to the failure. Since each motor includes 3 Hall sensorsand 4 poles, there are 12 Hall state changes per revolution of eachmotor. Remarkably, by monitoring the Hall state changes, resolution canbe increased by a factor of 12 such that a resolution of about 0.017degrees is provided at the output shaft of the actuator. In anembodiment, a corresponding movement at the top of the stick in FIG. 1can be about 0.0039 inch.

A total power failure of the helicopter's electrical power system cancause the actuators to lock in position for about five seconds using adynamic braking feature that is described in the above incorporated U.S.patent application Ser. No. 13/763,574 (Attorney Docket No. HTK-2). Thisfive second time period is generally more than sufficient for the pilotto take over control. In this regard, regulatory requirements mandate atime period of only three seconds. In this regard, the autopilot doesnot let the cyclic stick flop over by releasing control responsive to apower failure. Even with both actuators locked, the pilot can stillperform control over the helicopter since there are override or forcelimited links 300 a (pitch, seen in FIG. 1) and 300 b (roll, seen inFIGS. 1 and 2) between each actuator and the cyclic stick. These linksare rigid for forces below an unseating value and compliant at higherforces to allow the pilot to safely maneuver and land the helicoptereven if disengagement of the system cannot be achieved. It has beenempirically demonstrated that a pilot can control the helicopter,including hovering and landing, with both actuators in what is referredto as a “locked” state. The locked state is provided by shorting allwindings of the actuator motors and is used in the dynamic brakingembodiment described above. The override links are described in detailbelow and may be referred to interchangeably as force-limited links. Ina helicopter that does not utilize a hydraulic interface to the cyclic,cyclic vibration isolators 302 a (pitch) and 302 b (roll) can be locatedon the output shaft of each actuator. The vibration isolators may beoptional for use with a helicopter having hydraulic assistance on thecyclic control since the hydraulic system generally provides damping ofcyclic oscillations. The vibration isolators reduce the two perrevolution oscillating motion, that is present in the R22 rotorcraftcontrol linkage and other light helicopters, to prevent vibratory loadson the rotorcraft control and to increase the fatigue life of theactuator components. The cyclic vibration isolators are described indetail below.

Having described the mechanical components of the autopilot in detailabove, it is now appropriate to describe the autopilot in terms of therelationship between the aforedescribed components and related controlelectronics. In particular, FIG. 5 is a block diagram of an embodimentof autopilot 12. In this regard, main unit 30 comprising enclosure 31,the pitch and roll actuators 60 and electronics 66 may be referred tohereinafter as the Motor Control Processor Unit (MCPU) or main autopilotunit 30. The MCPU includes three microprocessors, each of which may bereferred to as a Motor Control Processor (MCP). There are three MCPs,individually designated as MCP A, MCP B and MCP C. These processor unitseach access a sensor suite of tri-axial MEMS rate sensors and tri-axialMEMS accelerometers. The MCPs are used to provide an inner loop of anoverall control system having an inner control loop and an outer controlloop. The MCPs provide commands to brushless DC motors, Motor A andMotor B of pitch actuator 60 a and roll actuator 60 b, driving thecontrol system for the helicopter. All inter-processor communication canbe through a serial bus that is natively supplied on each of theprocessors. Data integrity can be protected, for example, through theuse of a cyclic redundancy check (CRC) incorporated into the datastream.

The Federal Aviation Administration (FAA) certifies airborne systemsoftware under a version of DO-178. At the time of this writing, DO-178Chas been released. This document specifies Design Assurance Levels(DALs) based on the criticality of software failure in a given system.For example, DAL A is designated as “catastrophic” and is assigned wherea failure may cause a crash. As another example, DAL C is designated as“major” and is assigned where a failure is significant and may lead topassenger discomfort or increased crew workload. Each one of the threeMCPs can execute identical DAL A software to constitute atriple-redundant system. The motor control processors are interconnectedso that they can share data. Each processor reads its sensor suite andcompares its data with sensor data coming from the other two processorsfor purposes of consistency and each motor control processor computesaverages of all the corresponding sensors to use for further processing.In another embodiment, median values can be determined, as opposed toaverages. Sensor data determined to be erroneous is eliminated from theaveraging. A warning signal of sound and/or light can be sent toautopilot display processor unit (ADPU) 28 on instrument panel 20 (FIG.1). Haptic feedback such as, for example, stick shaking can be usedalone or in combination with other warning signal indications. In anembodiment, status lights, best seen in the enlarged inset view of theADPU in FIG. 1, include green (normal), amber (caution) and red(critical failure), as well as dual warning horns to provide systemstatus indications. The warning horns also provide system statusnotifications and alarms along with the status lights. Both the statuslights and horns interface directly to the MCPs. In some embodiments,sounds and/or warnings can be transmitted over the helicopter audiosystem such that notifications can be heard in the pilot's headset aswell as or in lieu of being issued from the horn. Complementing thestatus lights and horns is a display which provides current autopilotsystem settings such as engagement status, track, slaved gyroscopicheading, altitude, speed over ground and any warning messages. Also onthe panel is a testing button which initiates an Initiated Built-In Test(IBIT).

The MCPs also read Hall sensor data from the actuator motors, which canbe used to indicate the current position of each actuator, and a commandsignal coming from an autopilot display processor (ADP) which forms partof the ADPU. In this regard, the ADPU serves as the outer control loopto provide command signals to the inner loop. Using all these data, eachMCP calculates a motor control signal for the motors in terms of a PWM(Pulse Width Modulation) and direction of rotation. Each processor alsouses the Hall sensor data to control the power connections to thearmature of the brushless motors assigned to it. Each MCP compares itsPWM command signal and rotation direction for the pitch and rollactuators with commands generated by the other two MCPs for agreement.Since all processors are using the same data to compute motor controlsignals, they should produce identical output signals. Signals foragreement/disagreement with the other two processors are sent to avoting section 200, an embodiment of which is shown in detail in theabove incorporated U.S. patent application Ser. No. 13/763,574 (AttorneyDocket No. HTK-2), but the operation of which is described in furtherdetail below for purposes of completeness. In addition to votehandling/arbitration, section 200 also serves as a pass through for Hallsensor data from each of the motors to an appropriate one of the MCPs.As discussed above, the Hall sensor readings are used to generate motorcontrol signals for the brushless DC motors and can serve to provide ahigh resolution indication of the output shaft position of the actuator.

As described above, each actuator includes motor A and motor B. Eachindividual motor is controlled by one MCP. Thus only MCP A and MCP Bcontrol motors. In particular, MCP A controls motor A in each of pitchactuator 60 a and roll actuator 60 b, while MCP B controls motor B ineach of pitch actuator 60 a and roll actuator 60 b. MCP C (the thirdprocessor) does not control a motor but performs all calculations togenerate stick commands as if it were controlling a motor. In thisregard, a third motor can readily be added to each actuator (see FIG. 4)that would engage gear 102 in the same manner as motor A and motor B,but responsive to MCP C. The latter, however, votes in a manner that isidentical to the other two processors. For example, if MCP A and MCP Cagree on the control of the pitch motor, but MCP B does not, then MCP Bwill be voted out from control of its pitch motor, MCP B will stillcontrol its roll motor unless MCP A and MCP C also vote out control ofthat motor. On the other hand, if MCP C is voted out, no actuator motorswill be affected, but a warning light and horn can be actuated as wouldbe the case for the MCPs which control motors.

The actuators are designed such that either one of motor A or motor B isindependently capable of driving the actuator to control the helicopter.The output shaft of a failed motor will be rotated by the remainingmotor. If one of MCP A or MCP B is voted out, the autopilot can continueto function despite the fact that each of these MCPs controls motors. Asstated, there can be a warning light and a brief sounding of the horn tonotify the pilot that there has been a non-critical autopilotmalfunction.

The MCPs have full authority over the controls and are rate limited to asuitable value such as, for example, 5 inches per second. The MCPcontrol section is the only portion of the autopilot that can create acritical or major hazard malfunction. Accordingly, the MCPU is designedas triple-redundant with DAL A designated software for purposes ofoperating the inner loop of the autopilot. These factors greatly reducethe probability of a critical failure. Applicants recognize, however,that the software corresponding to the outer loop can be partitionedfrom the inner loop software in a way that allows the outer loopsoftware to be designated at a lower DAL C certification.

The outer loop software is handled by the ADP (Autopilot DisplayProcessor) in ADPU 28. The MCPs convert requested autopilot commandsfrom the ADP into actuator control signals that can drive the actuatormotors within defined operational limits. In this regard, it should beappreciated that DAL A software is handled by the triple redundant MCPswhile DAL C, outer loop software is handled by a completely differentprocessor. By way of still further explanation, a single executable runson each MCP. The MCPs, which may be referred to as the triplexprocessors, can execute identical software. Thus, the autopilot controllaws are partitioned between the ADP and triplex processors. The ADPprocesses the outer loop dynamics and autopilot modes while the triplexMCPs process the inner loop dynamics. The ADP further provides thepilot's graphical and testing interface to the autopilot and executesthe autopilot control laws to determine actuator commands based onsensor and GPS data. Accordingly, the ADP interfaces directly with theGPS and triaxial magnetometers and indirectly with triaxialaccelerometers and triaxial rate gyros of the MCPs which provide theroll-pitch attitude, position, altitude, ground speed, course andheading data. The ADP monitors the health of these sensors but does notcheck the validity of the data. The IBIT test switch also interfaces tothe ADP. In another embodiment, the ADP can be designed in the samemanner as the MCPU with triple redundancy. With both the MCPU and ADP ina triple redundancy configuration, the autopilot can tolerate a singlefailure in either of these units and still remain fully functional.

The MCPs accept data from the ADP which can include commands as well asdata from an external GPS. The data can be screened by each MCP todetect errors or malfunctions. The control command is rate-displacementlimited by the MCPs. The MCPs will not allow a command from the ADP tocreate a hazardous response from the helicopter. GPS data is used by theADP. The GPS and magnetometer data are both used in the MCPs to removedrift errors associated with the rate sensors of each sensor suite andto determine roll, pitch and heading. The GPS data can also be checkedfor errors.

The MCPs constantly monitor for both internal and external faults. Inthe event of an ADP failure, any one MCP can immediately recognize thesituation based on update rate and control signal conformity. Inresponse, the MCPU, in one embodiment, will then cause the inner controlloop to hold the helicopter straight and level. In another embodiment,the MCPU can act in the manner of a SAS (Stability Augmentation System)or a dead reckoning system and control the helicopter based on internalrate signals. The MCPs will attempt to hold zero rates and/or headingand also actuate a horn and light to indicate a failure. It has beenempirically demonstrated that the helicopter can maintain prolongedflight with only MCP control, providing more than ample time for thepilot to take control and disengage the autopilot. The ability to detectexcessive autopilot response resides in the triplex motor controllers asdetailed herein. The triplex processors monitor sensors and also checkto confirm that calculated responses are within limits. Pitch and rollcommands from the ADP are limited based on such command filtering byeach of the triplex processors. Each triplex processor can detectwhether a limit has been exceeded and can initiate safe shut down of theautopilot. Pitch and roll axis commands can be monitored identically butwith different limit values. The monitors are dynamic; that is, thelimit values can be frequency/rate dependent. Redundancy managementfeatures for each axis can include stick rate limiting and body ratemonitoring.

The sensor suite of each MCP can also include memory such as, forexample, EEPROM or other suitable memory. If there is an error detectedby an MCP during operation, the error code can be stored in the EEPROMof the sensor suite associated with the MCP. The EEPROM can later beread in the context of determining the cause of failure. The EEPROMs canalso contain parameters specific to the model of the helicopter in whichthe autopilot is installed such as, for example, control loop constants,sensor offsets and gains.

Referring to FIGS. 6 and 7 in conjunction with FIGS. 1-3, it should beappreciated that enclosure 31 can provide benefits with respect toinstallation of the autopilot system into a helicopter that areheretofore unseen. FIGS. 6 and 7 provide perspective views of anembodiment of the enclosure including main electronics unit 66 andactuators 60 a and 60 b mounted within the enclosure such that outputshafts 142 a and 142 b of each respective actuator extends outward fromthe interior of the enclosure. It is noted that actuator shaft 142 a islonger than actuator shaft 142 b based on installation specificrequirements. The length of each actuator shaft can be customized inview of an installation in a given type of helicopter. FIG. 6 provides afront/top, perspective view, while FIG. 7 provides a back/bottom,perspective view. Each actuator housing 82 (FIG. 3) can be configured toreceive suitable fasteners 402 such as, for example, threaded fastenersthrough holes that are defined by enclosure 31 in order to support theactuator within enclosure 31. FIG. 6 illustrates first pairs offasteners 402 that extend through the top cover of the enclosure andinto each actuator housing while FIG. 7 illustrates second pairs offasteners 402 that extend through the top cover of the enclosure andinto each actuator housing. Thus, the actuators and main electronicsunit are commonly received within the interior space defined by theenclosure and mounted/captured against interior surfaces of the walls ofthe enclosure.

The main electronics unit at least includes motor drivers for drivingmotors (BLDC motors in the present example) as well as the inner controlloop of the autopilot system as shown, for example, in FIG. 5.Additional components that can be received within enclosure 31 caninclude, for example, a power supply section for powering the entireautopilot system. Each MCP processor can be provided with an independentpower supply that is commonly housed within enclosure 31. While anL-shaped enclosure is illustrated for installation in the R22, it shouldbe appreciated that the enclosure can be of any suitable shape in viewof an intended application and is not limited to L-shaped so long as themain electronics unit and actuators can be commonly received within theenclosure. Enclosure 31 can be configured to handle the forces generatedby the actuators with the actuators mounted directly on the walls orplates of the enclosure, for example, using fasteners. In the presentembodiment, each actuator is secured against the bottom plate, the topplate and one of the plates through which the actuator shafts pass. Theuse of enclosure 31 for purposes of housing the bulk of the electronics,control system and actuators of the autopilot embodies a minimallyintrusive lightweight package. The overall weight of the componentssupported by enclosure 31 and the enclosure itself can be less than 8pounds with the weight of the enclosure itself being less than 7 pounds.In contrast, prior art autopilot systems of which Applicants are awarerequire separate installation of the main electronics assembly and eachactuator. Typically, each actuator, in a prior art autopilot, isindependently and directly mounted to the structure of the helicopteritself with the need for special or customized mounting provisions suchas structural reinforcements associated with each actuator. The use ofenclosure 31, as taught herein, avoids the need for such complexindependent installation associated with the actuators to providebenefits that can include reducing installation time as well as reducingthe overall weight of the autopilot system. In fact, the installation ofenclosure 31 can be accomplished through the use of a straightforwarddrill template, resulting in a great degree of accuracy insofar as thepositioning of the enclosure while being highly economical.

Turning to FIG. 8, an embodiment of enclosure 31 is shown in aperspective view, generally from above, without components installedtherein and with selected elements rendered as transparent to reveal theappearance of otherwise hidden elements. The enclosure can be formed, byway of non-limiting example, by a top plate 600 a bottom, mounting plate602, front covers/plates 604 and back covers/plates 606. The back coversdefine openings 610 for receiving output shafts of the actuators. Thevarious panels or plates of the enclosure can be attached to one anotherin any suitable manner such as, for example, using welding and orsuitable fasteners such as rivets and/or threaded fasteners. In thepresent embodiment, mounting plate 602 includes front flanges 614 thatextend upwardly. Front covers 604 can include top flanges 616 that canextend over top plate 600. Top plate 600 can include front flanges 620that extend downward in the view of the figure. Fasteners 622, one ofwhich is shown, can be received in appropriate openings for purposes ofsecuring the enclosure components to one another. In the presentembodiment, the actuators can be supported by multiple panels of theenclosure such as the rear sidewalls, top plate 600 and mounting plate602. In this regard, actuator housings 82 (FIG. 3) can contribute to theoverall structural rigidity of enclosure 31.

Referring to FIG. 2, the actuator shafts can interface to the controllinkage components through sidewalls or bulkheads of the helicopterbaggage compartment such as, for example, a first bulkhead 630 throughwhich the pitch actuator output shaft passes and a second bulkhead 632through which the roll actuator shaft passes. Bulkhead 630 has beenrendered as transparent for illustrative purposes. As will be furtherdescribed and by way of non-limiting example, enclosure 31 can beremovably fixedly mounted against bulkheads 630 and 632. In this regard,it is noted that a clearance is present below the enclosure in thepresent embodiment since the floor of the baggage compartment below theunit is not flat. Of course, such details can be dependent oninstallation in a given helicopter and are not intended as limiting.Fasteners 634, as seen in FIG. 7 and two of which are partially visiblein FIG. 2, can be used for purposes of mounting enclosure 31 tobulkheads 630 and 632 by installing these fasteners through therespective bulkheads to threadingly engage the actuator housings.Additional fasteners can be used to secure enclosure 31 to thebulkheads, as will be described. In any installation, main enclosure 31can provide both structural support and shielding. The enclosure can beformed from any suitable material such as sheet materials of aluminum.

FIGS. 9 and 10 are perspective views looking up toward an embodiment ofenclosure 31 in an installed condition from somewhat different anglesbut with both figures limited to showing enclosure 31 having pitchcontrol linkage 32 a and roll control linkage 32 b installed andinterfaced to a lower end of stick 14, which is only partially shown.The lower end of the stick can move in any lateral direction responsiveto movement induced by the pilot. A friction arrangement 680 can beprovided as original equipment in the helicopter and is adjustable usinga tension knob 682 in a manner that will be familiar to those ofordinary skill in the art. The distal end of force-limited link 300 a(pitch) includes a pivot type mount 700 a such as, for example, a balland socket type mount that is pivotally attached to the lowermost end ofstick 14. For roll autopilot actuations, a pivot type mount 700 b, at adistal end of force-limited link 300 b, can be of the same type and ispivotally attached to the lowermost end of stick 14 via a roll linkagebar 704. The latter is itself configured for pivotal attachment to thelowermost end of the stick. Original torque tube 25 a and roll controlrod 25 b are understood to be present, as seen in FIG. 1, but have notbeen shown in the present figure for purposes of illustrative clarity.The torque tube and roll control rod can be unchanged from theiroriginal forms for purposes of communicating pilot actuations to theswash plate of the helicopter in order to accomplish cyclic control.

Still referring to FIGS. 9 and 10, output shaft 142 a (see FIG. 2) ofpitch actuator 60 a is received by vibration isolator 302 a which is inturn pivotally attached to force-limited link 300 a. Similarly, outputshaft 142 b (see FIG. 2) of pitch actuator 60 b is received by vibrationisolator 302 b which is in turn pivotally attached to force limited link300 b. As noted above, actuator shaft 142 a is longer than actuatorshaft 142 b. As best seen in FIG. 10, a bearing assembly 708 can be usedto support the extended actuator shaft. The bearing assembly can beattached to enclosure 31 in any suitable manner such as, for example, byusing removable fasteners 709, as illustrated. In the presentembodiment, bulkhead 630 (seen in FIG. 2) is captured between thebearing housing and the enclosure with fasteners 709 threadinglyengaging the housing of pitch actuator 60 a. Enclosure 31, theactuators, the vibration isolators and the force-limited links can bepreassembled for installation as a unit into the helicopter. It shouldbe appreciated that a bolt forming a pitch/roll attachment point 710 canbe extended for receiving the distal end pivot mount of tension link 300a (pitch) as well as the pivot mount of roll linkage bar 704. Theopposite end of roll linkage bar 704 is pivotally attached to the distalend of roll force-limited link 300 b which also defines an opening forattachment of roll friction arms 714. An access hole 716 (FIG. 2) allowsthe roll actuator control link to pass through from the landing geartunnel to the console tunnel of the helicopter. Enclosure 31 can befurther removably secured against bulkheads 630 and 632, for example,using threaded fasteners 720, only two of which are individuallydesignated, that can extend through respective ones of the bulkheads toengage nut-plates of any suitable type mounted within the interior ofenclosure 31. It is noted that openings 722 have been shown that arelikewise configured with nut-plates for receiving additional fasteners720. Any suitable number of such fasteners can be used. Moreover, theuse of fasteners 720 is not intended as limiting and the enclosure canbe secured into an installed position in any suitable manner.Removability of the enclosure can facilitate the configuration ofoverall main autopilot unit 30 as a non user serviceable item. That is,if any concern arises with respect to the operation of the mainautopilot unit, it can readily be completely removed from the helicopterand delivered to a service facility for repair.

As seen in FIGS. 9 and 10 and aside from enclosure 31, the autopilotmechanical system includes three major mechanical components: actuators60, vibration isolators 302 and force-limited links 300. The actuatorsimpart motion to the rotorcraft control system while the vibrationisolators reduce the two per revolution oscillating motion that ispresent, for example, in a light rotorcraft control linkage such as theR22, to prevent vibratory loads on the rotorcraft control and toincrease the fatigue life of the actuator component, as discussed above.Force-limited links 300 transmit motion from the actuator to therotorcraft control linkage while allowing the pilot to override inputsfrom actuators 60. In a helicopter having a control linkage that is notsubject to cyclic oscillation such as, for example, a helicopter havinga hydraulically assisted cyclic control system, vibration isolators 302can be replaced with rigid arm members that can be non-resilient atleast from a practical standpoint.

Having described actuators 60 in detail above, attention is now directedto FIG. 11 which illustrates additional details with respect tovibration isolator 302. As discussed above, vibration isolators 302 aand 302 b can be of identical construction and therefore the referencenumber 302 can refer to either vibration isolator. The vibrationisolator includes weight arms 1100 a and 1100 b which can be identicalin construction. The weight arms can be referred to collectively by thereference number 1100 and can be formed of any suitable sheet materialsuch as, for example, aluminum. A first end of the weight arms receive apin 1102 that is configured for engaging one end of force limited link300 (FIGS. 9 and 10). The weight arms can be fixedly attached to pin1102 in any suitable manner such as, for example, using a pressed fit orwelding. In this regard, it is not necessary for the pin to pivotallyengage the weight arms. Pin 1102 can define a shoulder 1104 and includea threaded distal end such that one end of the force limited link can becaptured between shoulder 1104 and a nut 1108. Opposite ends of weightarms 1100 are fixedly attached to a tuning weight 1110 that can beformed from any suitable material such as, for example, brass.Attachment can be accomplished in any suitable manner such as throughthe use of threaded fasteners 1112. Each weight arm can further includea spring keeper tab 1114 a or 1114 b that can be integrally formed andbent transversely to the main body of the weight arm. The spring keepertabs will be described in further detail below. The weight arms receivea bearing set 1120 having a pin 1122 that extends through a pivot end ofa control arm 1130 such that the weight arms can pivot in unison inrelation to control arm 1130. It is noted that each actuator shaft 142(see FIGS. 7 and 9) supports one control arm, as shown in FIG. 9 anddesignated by the reference numbers 1130 a and 1130 b.

Referring to FIG. 12 in conjunction with FIG. 11, the former is aperspective, exploded view which illustrates one control arm 1130 andrelated components. Each control arm defines a force limited linkaperture 1132 and an actuator shaft aperture 1140 that is configured toreceive one of actuator shafts 142. When installed, a saddle 1142 on theactuator shaft is aligned with a key aperture 1144 that receives across-pin key 1148. The latter includes through openings 1150, forexample, to receive safety wire (not shown) to secure the cross-pin keyin its installed position. Each control arm can define opposing pockets1152 of removed material which, as an option, can reduce the weight ofthe control arm.

Attention is now directed to FIGS. 11-13. FIG. 13 is a perspective viewof vibration isolator 302 with weight arm 1100 a rendered as transparentfor purposes of illustrative clarity. A first isolation spring 1302 a iscaptured between keeper tab 1114 a (FIG. 11) of weight arm 1100 a andcontrol arm 1130 while a second isolation spring 1302 b is capturedbetween keeper tab 1114 b of weight arm 1100 b and control arm 1130.Springs 1302 can be formed from any suitable material such as, forexample, steel or corrosion resistant alloys such as nickel chromiumbased alloys. A spring keeper 1310 passes through the interior space ofeach isolator spring and through control arm 1130 such that the springkeeper is captured between keeper tabs 1114 a and 1114 b. In thisregard, the spring keeper can include, for example, a rectangular orother suitable cross-sectional shape with a head 1312 of reduced size incomparison to the cross-sectional shape of the overall length of thespring keeper. Thus, the base of head 1312 defines one or more shouldersthat can be received against an inside surface of each keeper tab withhead 1312 itself received in a suitable/complementary opening defined bythe keeper tab as shown in FIG. 11. Control arm 1130 (FIG. 12) definesopposing spring pockets 1320 that receive respective inner ends of theisolation springs against a shoulder 1324 while the outer end of eachisolation spring biases against a spacer 1328 that, in turn, biasesagainst one of the keeper tabs. It is noted that one of the spacers hasbeen rendered as transparent in the view of FIG. 13 in order tofacilitate a view of one end of spring keeper 1310. In an embodiment,each shoulder 1324 can be angled such that when an associated isolationspring 1302 is fully compressed, the shoulder defines a plane that is atleast approximately parallel to the associated keeper tab in aconfronting relationship therewith.

Vibration isolator 302 serves as a variable compliance device actingbetween one actuator 60 and an associated force-limited link 300. Whilefurther detail will be provided below, for the moment is sufficient tonote that force-limited links 300 are effectively rigid with respect tothe cyclic vibration or stick shaking frequency and are likewiseeffectively rigid responsive to normal autopilot actuations fromactuators 60. The vibration isolator is very stiff for low frequencyinputs provided from actuator 60 via the actuator shaft but is verycompliant at the two per revolution cyclic vibration frequency of 17.5Hz that is present on force limited link 300. During operation, with theactuator shaft at a fixed position, weight arms 1100 oscillate aboutpivot 1122 in a way that alternately compresses isolation springs 1302between each weight arm keeper tab 1114 and each control arm 1130responsive to cyclic stick shaking. Thus, the actuator shaft iseffectively isolated from the two per revolution cyclic vibrationfrequency. At the same time, the frequency response of each actuator 60is approximately 3 Hz for purposes of rotating the actuator shaft toprovide autopilot actuations. Accordingly, there is sufficientseparation between the maximum control frequency of actuator inputs andthe two per revolution cyclic vibration frequency such that controlforces from the actuator output shaft pass through the vibrationisolator and onto the helicopter cyclic control system via force tensionlink 300 while the rotor induced vibratory forces are isolated from theactuators. While the present embodiment of the vibration isolator isconfigured for use in the R22 helicopter having a cyclic vibrationfrequency of approximately 17.5 Hz, the vibration isolator can readilybe modified for operation with a helicopter that exhibits a differentcyclic vibration frequency.

Having described the vibration isolator of the present disclosure indetail above, additional details are now provided with respect to anembodiment that is configured for the 17.5 Hz cyclic vibration that istypical of a Robinson R22 helicopter. That is, the weight arms, servingas an overall output arm of the isolator, the tuning weight andassociated dimensions make up a resonant system having a resonantfrequency that is at least approximately equal to the cyclic vibrationfrequency of the helicopter. Of course, the tuning weight can beadjusted to accommodate a range of different resonant frequencies.Similarly, various dimensions can be adjusted to change the resonantfrequency, as desired. In the present embodiment, springs 1302 have acombined spring constant of 16.6 lbf/in that is located 2.0 inches frompivot 1122. The tuning weight has an equivalent mass of 0.2 lbm(pound-mass) located 3.3 inches from pivot 1122.

Attention is now directed to FIG. 14 which is a perspective view of anembodiment of force-limited link 300. Initially, it should beappreciated that the force-limited link is designed to serve as a rigidpush-pull rod at loads below a predetermined break-away force and to becompliant at loads above the break-away force. This allows the pilot toreadily override actuators 60 at any time, even if one or both actuatorsare jammed. The design can provide for override forces that are greaterthan normal control forces but which override forces are easily managedby the pilot. In the present embodiment, a first end 1402 of the forcelimited link is designed for mounting upon shaft 1102 of FIG. 11 whilesecond, opposing end 700 is designed for mounting to engage the cycliccontrol system as shown, for example, in FIGS. 10 and 11 wherein pivotmount 700 a of pitch force-limited link 300 a (FIGS. 9 and 10) isattached to the bottom of the stick and pivot mount 700 b of roll forcelimited link 300 b is attached to roll linkage bar 704. Each end of theforce limited link can utilize a ball and socket type mount which canaccommodate motions such as arcuate and side-to-side motions that arecharacteristic of the movement of the bottom of stick 14 under thecontrol of the pilot. First end 1402 can include an end cap 1404 andlock ring 1406 that threadingly engage a housing 1410. The latter candefine an adjustment thread 1412. Accordingly, the length of a mainportion of the internal cavity is defined by the end cap in cooperationwith the housing and can therefore be adjusted, as will be furtherdescribed. It is noted that the orientation of the force limited linkcan be reversed end-for-end in a given installation so long asclearances are adequate.

Referring to FIGS. 15 and 16 in conjunction with FIG. 14, FIG. 15 is aperspective exploded view of force-limited link 300 shown here toillustrate details with respect to its internal components while FIG. 16is an assembled partially cutaway perspective view. Housing 1410supports a shaft 1412 for linear movement 1414 as illustrated by adouble-headed arrow. Shaft 1412 defines a shoulder 1418 against which aspring bias disk 1420 can be received. A reduced diameter end portion1422 of the shaft extends from shoulder 1418 to a spring bias head 1424.Spring bias disk 1420 is configured for sliding/lateral movement alongend portion 1422 in the direction of arrow 1414 responsive to externalbiasing forces. When assembled, a top hat 1430 internally receivesspring bias head 1424 for movement according to arrow 1414. A crown 1432of the top hat is, in turn, received within an interior space defined bya spring 1440 with the spring in a preloaded state. Like the spring biasdisk, a crown end 1442 of the top hat is configured for lateral/slidingengagement along the length of end portion 1422 of shaft 1412. Spring1440, for example, is a helical coil spring that can be formed using asuitable material that can be corrosion resistant. Top hat 1430 andspring bias disk 1420 can be formed from any suitable material such as,for example, aluminum. Shaft 1412 can be formed from any suitablematerial such as, for example, stainless steel. Housing 1410 and end cap1404 can be formed from any suitable material such as, for example,aluminum.

FIG. 16 illustrates the force-limited link in what can be referred to asa relaxed state with spring 1440 captured between a brim 1444 of the tophat and spring bias disk 1420 such that the spring is preloadedtherebetween. In this relaxed state, bias head 1424 is resilientlybiased against a crown end 1442 of the top hat. Adjustment ring 1406 canbe adjusted to compensate for tolerances between the various componentssuch that spring 1440 is extended as fully as possible while stilltaking up any play between the various components. That is, brim 1444just contacts end cap 1404 while spring bias disk 1420 just contacts ashoulder 1450 of housing 1410 and spring bias head 1424 is resilientlybiased against crown end 1442 of the top hat. As described above, therelaxed state establishes an amount of compression on spring 1440 thatis referred to as a spring preload. The preload establishes theunseating force and represents a minimum level of compression of thespring in the assembly such that it, therefore, will begin to compresswhenever external forces are applied to the force limited link thatexceed the spring preload/unseating value. In an embodiment, spring 1440can be subjected to approximately one-half of its ultimate deflectedload as the preload value. In the present embodiment, the spring preloadis selected as at least approximately 25 lbs. A suitable range ofpreload values can be selected to accommodate the requirements of aparticular helicopter.

FIG. 17 is a partially cutaway perspective view that illustrates a fullyretracted state of force-limited link 300. As shown, shaft 1412 isreceived to a maximum extent within housing 1410 having bias head 1424and brim 1444 received against end cap 1404. At the same time, spring1440 is subjected to a maximum level of compression in the assembly.FIG. 17 demonstrates subjecting spring 1440 to an external compressiveforce of sufficient magnitude above the spring preload to exceed theunseating force and move the ends of the force-limited link toward oneanother, thereby reducing the length between its opposing ends. Ideally,the force-limited link unseats at a preset load and exerts a uniformforce throughout its travel. From a practical standpoint, the spring isselected to provide the required preload force and not exceed amanageable force for the pilot when the link is at full deflection.

FIG. 18 is a partially cutaway perspective view that illustrates a fullyextended state of force-limited link 300. As shown, the force limitedlink is subjected to an external force that is of sufficient magnitudeto cause a full extension of the link. It is noted that this force canbe of at least approximately the same magnitude as the force, althoughopposite in effective direction, which produces the fully retractedstate of FIG. 17. That is, although spring 1440 is laterally displacedin FIG. 18 as compared to FIG. 17, the spring is compressed to the samelength in both figures. Assuming for descriptive purposes that first end1402 is fixed in position, the external force pulls on second end 700such that crown end 1442 of the top hat is pulled to the right in theview of the figure in a way that causes brim 1444 to compress spring1440. This movement results in displacing reduced diameter end portion1422 of shaft 1412 to the right in the view of the figure until biashead 1424 forces crown end 1442 to contact spring bias disk 1420. In anembodiment, the spring constant of spring 1440 should be sufficientlylow to allow the assembly to reach its full ends of travel as shown inFIGS. 17 and 18. While the force-limited link has been illustrated inthree different operational conditions, one of ordinary skill in the artwill appreciate that the device transitions from one state to another ina manner that is consistent with these descriptions.

The foregoing description of the invention has been presented forpurposes of illustration and description. It is not intended to beexhaustive or to limit the invention to the precise form or formsdisclosed, and other modifications and variations may be possible inlight of the above teachings wherein those of skill in the art willrecognize certain modifications, permutations, additions andsub-combinations thereof.

What is claimed is:
 1. In an autopilot system that is configured forautomated control of a helicopter by actuating a cyclic control of thehelicopter which cyclic control is subject to a cyclic vibrationfrequency responsive to rotation of a rotor of the helicopter, anapparatus comprising: a cyclic vibration isolator that is removablysupported by said actuator shaft for co-rotation therewith and coupledto the cyclic control of the helicopter to exhibit a resonant frequencythat is at least approximately matched to the cyclic vibration frequencyfor movement of the cyclic control relative to the actuator shaft suchthat the cyclic vibration frequency is attenuated at the actuator shaftand output rotations of the actuator shaft below the resonant frequencyare coupled to the cyclic control.
 2. In an autopilot system that isconfigured for automated control of a helicopter by actuating a cycliccontrol of the helicopter which cyclic control is subject to a cyclicvibration responsive to rotation of a rotor of the helicopter, a cyclicvibration isolator comprising: a control arm that is removably attachedto the actuator shaft to support the cyclic vibration isolator such thatthe control arm co-rotates with the actuator shaft; an output arm thatis coupled to the cyclic control of the helicopter and therefore subjectto said cyclic vibration; and a resilient arrangement that is capturedbetween the control arm and the output arm such that the output armoscillates responsive to the cyclic vibration and relative to thecontrol arm to mechanically isolate the actuator shaft from the cyclicvibration while transferring rotational actuation motions of theactuator shaft to the output arm to thereby transfer the rotationalactuation motions to the cyclic control for autopilot actuation of thecyclic control.
 3. The cyclic vibration isolator of claim 2 wherein thecontrol arm, the output arm and the resilient arrangement are configuredto cooperate such that the output arm oscillates at a resonant frequencythat is at least approximately equal to a cyclic vibration frequency ofthe cyclic vibration.
 4. The cyclic vibration isolator of claim 2wherein the output arm pivots about a pivot axis that is defined by thecontrol arm and the pivot axis is spaced apart from the actuator shaft.5. The cyclic vibration isolator of claim 4 wherein the output armsupports an output shaft that is coupled to the cyclic control and theoutput shaft is on an opposite side of the pivot axis with respect tothe resilient arrangement.
 6. The cyclic vibration isolator of claim 5wherein the output arm supports a tuning weight on an opposite side ofthe pivot axis with respect to the output shaft.
 7. The cyclic vibrationisolator of claim 4 wherein the output arm includes an output shaft thatis coupled to the cyclic control and the output shaft is on an oppositeside of the actuator shaft with respect to a position at which theresilient arrangement contacts the control arm.
 8. In an autopilotsystem that is configured for automated control of a helicopter bydriving an actuator having an actuator shaft to actuate a cyclic controlof the helicopter, an apparatus comprising: a force limited link havinga first end, a second end and a variable length therebetween orientedalong an elongation axis with the first end coupled to the actuatorshaft and the second end coupled to the cyclic control, said forcelimited link having a relaxed length as the variable length between thefirst and second ends when less than an unseating force is applied tothe first and second ends along said elongation axis to providecompliant movement of the cyclic control responsive to the actuator andconfigured such that the variable length changes from the relaxed lengthresponsive to an external force that is applied to the first and secondends along the elongation axis which is equal to or greater than theunseating force to provide for a pilot override of actuator cycliccontrol.
 9. The apparatus of claim 8 wherein the variable lengthdecreases from the relaxed length responsive to the external force beingapplied in compression along the elongation axis and the variable lengthincreases from the relaxed length responsive to the external force beingapplied in tension along the elongation axis.
 10. The apparatus of claim9 wherein the first end is supported by a housing that defines a housinginterior and an extension shaft is supported to extend outward from thehousing interior and for movement along the elongation axis and a distalend of the extension shaft supports the second end.
 11. The apparatus ofclaim 10 wherein the housing interior receives a resilient member thatis in a minimum compressed condition for said relaxed length andresponsive to movement changing the relaxed length the resilient memberis subjected to an additional compression.
 12. The apparatus of claim 11wherein the resilient member is a helical coil spring having first andsecond opposing ends.
 13. The apparatus of claim 12 wherein the helicalcoil spring is supported such that the first end of the spring movestoward the second end of the spring when the variable length is extendedbeyond the relaxed length such that the spring is displaced toward thesecond end of the force limited link.
 14. The apparatus of claim 12wherein the helical coil spring is supported such that the second end ofthe spring moves toward the first end of the spring when the variablelength is reduced to less than the relaxed length such that the springis displaced toward the first end of the force limited link.
 15. In anautopilot system that is configured for automated control of ahelicopter by providing electrical command signals to an actuator toactuate a cyclic control of the helicopter which cyclic control issubject to a cyclic vibration responsive to rotation of a rotor of thehelicopter, an autopilot linkage comprising: a force limited link toprovide a pilot override of cyclic control by said actuator, said forcelimited link having a first end, a second end and a variable lengththerebetween oriented along an elongation axis with the first endcoupled to the cyclic control, said force limited link having a relaxedlength as the variable length between the first and second ends whenless than an unseating force is applied to the first and second endsalong said elongation axis to provide compliant movement of the secondend with the first end and configured such that the variable lengthchanges from the relaxed length responsive to an external force that isapplied to the first and second ends along the elongation axis which isequal to or greater than the unseating force for said pilot override;and a cyclic vibration isolator that is removably supported by saidactuator shaft for co-rotation therewith and connected to the second endof the force limited link such that said cyclic vibration is coupled tothe cyclic vibration isolator by the force limited link and the cyclicvibration isolator is configured to exhibit a resonant frequency that isat least approximately matched to a cyclic vibration frequency of thecyclic vibration to allow movement of the second end of the forcelimited link relative to the actuator shaft at the cyclic vibrationfrequency while isolating the actuator shaft from the cyclic vibrationfrequency and for transferring output rotations of the actuator shaftbelow the resonant frequency to the force limited link for transfer tothe cyclic control.